The present invention relates to a rocket drive system with air intake and including cyrogenic fuels for utilization as a space vehicle carrier system, as well as for use in a supersonic and hypersonic aircraft i.e. aerodynamic flight systems.
Presently, only those kind of rocket drives are used for space vehicle carrier systems which carry along all of the fuel, as well as an oxidation medium. However, as long as the carrier is still within the atmoshphere, one could use propulsion systems with air intakes just as they are used in supersonic and hypersonic aircraft, so as to take the necessary oxygen from the atmospheric environment as long as the rocket and propulsion system is, in fact, still in the atmosphere.
As stated, the presently known rocket drives and rocket propulsion systems and engines are independent from the presence or absence of air on the outside. For this reason they not only carry all of the fuel, but also all of the oxidizing medium necessary for combustion. Aside from this independence, there is a definite advantage in such a system which is launched from zero motion, zero speed condition, and that is its capabality of covering the entire speed and altitude range during any mission. A single mode propulsion ensures compact design, high thrust density, such as 20-200 N/cm.sup.2, and relatively low thrust specific mass, such as 1-2 kg/kN. On the other hand, all these advantages are, to a considerable extent offset by a large specific fuel consumption, such as 0.25 kg/kN.s. This fuel consumption is primarily determined by the presently preferred pairing of cryogenic fuel such as liquid hydrogen, LH.sub.2, with liquid oxygen LOX, whereby simply in terms of quantity 80% is occupied by the oxidizer. Cryogenic systems of this type use approximately 50% of the entire fuel (i.e. fuel proper plus oxidizer) for the accelerated ascent through the lower, relatively dense atmosphere, until attaining speeds in the range from 1500 to 1800 m/s, which is about Mach 5-6. The remaining fuel is used to accelerate the vehicle up until attaining orbital speeds, such as 7600 m/s.
This high fuel consumption for the accelerated ascent through the lower atmosphere inevitably leads to a relatively low and poor payload-to-fuel ratio, and an overall large mass and weight on launching. The high fuel consumption requires, moreover, the installation of propulsion systems which provide very high thrust, and in spite of the relatively low thrust specific mass, they are still heavy. Therefore, for reasons of operative simplicity, safety end lower operating cost, a single stage carrier system is pushing the border of, what is technilogocally feasible, bearing in mind that the risk of developing such system is very high. Another operative disadvantage of a pure rocket propulsion carrier system is the simple fact that the system itself cannot be transferred on its own between various locations on ground, for example, between the manufacturing plant and the launch site or between landing and launching, or landing point and manufacturing facility.
Propulsion drives with air intake, generally, will take the oxygen necessary for combustion from the environment as long as a sufficient amount is available. Consequently, their fuel consumption is very low, such as 0.02 kg/kN.s. Unfortunately, the known air intake systems are disadvantaged by large dimensions, relatively low thrust density, and relatively large thrust specific mass such as 10 kg/kN, while, of course, in addition, environmental air independent rocket facilities still have to be provided for to provide for the propulsion once the vehicle has left the air.
Among the known air intake type jet propulsions are, for example, the usual turbine jet propulsion systems, which are used in the high subsonic and moderate supersonic speed range. Conventional ram jets with subsonic combustion are used in the higher supersonic range, while the very high supersonic and hypersonic range uses ram jets with supersonic combustion and supersonic through-flow, also called SCRAMJETS. Finally, a multiple of so-called combination propulsions systems are known, such as turbo ram jets, turbo rocket drives, shrouded ram jet rocket drives, and so forth. In the so-called turbine air jet propulsion system, a turbine driven compressor causes air to be sucked into the combustion chamber. Such drives can be used right in the beginning for a takeoff from a standing position, with and without after burners. However, their thrust density is quite low and their thrust specific mass is much too large, at the present the known values are about 10 kg/kN. Moreover, the range of such turbojet drive and propulsion system is quite limited owing to the compression problems, so that, in fact, only a moderate to lower supersonic range can only be covered.
Ram jets have a diffuser which provides for the compression of captured air and feeds it to the combustion. Such jet propulsion system are incapable of providing adequate thrust from a standstill position. Moroever, they need a rocket assist drive or a turbojet drive just in order to provide for an initial accelleration, at least up to a speed in which the ram jet effect can be utilized. With increasing Mach number the ram temperature of the air limits the range of conventional ram jets with subsonic combustion to about Mach 7. Higher speeds require either a rocket assist drive or a transition to ram jet drive system with supersonic combustion. This type of propulsion system is known as SCRAMJET, and it avoids thermal problems resulting from the significant decelleration of air from hypersonic speed to subsonic speed. In addition, the significant static pressure losses in such a case are avoided. These SCRAMJETS avoid these problems specifically in that the hypersonic air flow is decellerated in the intake diffuser only to a moderate extent, down to a moderate supersonic speed. Realization and testing of a SCRAMJET drive, however, is an extremely expensive and risky development. Owing to the power limits of wind tunnels which limits are based on physical and technical considerations, it is extremely difficult to test any component in the hypersonic range.
Conventional ram jets with subsonic combustion, as well as SCRAMJETS, offer the advantage that within a quite large range of Mach number the specific fuel consumption remains low. Unfortunately, the thrust density is on the average also quite low, while the thrust specific mass is still relatively high. The drawback of the inability to accellerate from standstill is met in any other conceptualization by integrating these propulsion systems with a rocket drive which provides an initial acceleration until the ram jet or SCRAMJET operation can take over, while, finally, of course, the carrier system has to be accelerated to orbital speed by the rocket outside of the atmosphere. Particularly for single stage space vehicle carriers, one cannot, on the other hand, use different kinds of propulsion engines which work only part time and during certain phases, because any drive aggregate after its operating phase is just ballast which reduces the payload capacity of the carrier system.
Calculations have shown, for example, that a single stage carrier system with a cryogenic rocket drive and having been accellerated up to about 5-6 Mach, can, in fact, be brought into orbit, provided this drive aggregate which provided the initial accelleration will not consume more than 20% of the entire fuel that is needed, and if the additional components used for air intake operation, will not cause the thrust specific mass to exceed about 6 kg/kN.